Repeating airfoil tip strong pressure profile

ABSTRACT

A compressor section for a gas turbine engine includes a blade including a platform, a tip and an airfoil extending between the platform and tip. The airfoil includes a root portion adjacent to the platform, a midspan portion and a tip portion. Each of the root portion, midspan portion and tip portion define a meridional velocity at stage exit with the tip portion including a first meridional velocity greater than a second meridional velocity of the midspan portion. A blade for an axial compressor of a gas turbine engine and a method of operating a compressor section of a gas turbine engine are also disclosed.

CROSS REFERENCE TO RELATED APPLICATION

This application is a continuation of U.S. application Ser. No.15/255,663 filed Sep. 2, 2016.

STATEMENT REGARDING FEDERALLY SPONSORED RESEARCH OR DEVELOPMENT

The subject of this disclosure was made with government support underContract No.: NNC14CA36C awarded by NASA. The government therefore mayhave certain rights in the disclosed subject matter.

BACKGROUND

A gas turbine engine typically includes a fan section, a compressorsection, a combustor section and a turbine section. Air entering thecompressor section is compressed and delivered into the combustionsection where it is mixed with fuel and ignited to generate a high-speedexhaust gas flow. The high-speed exhaust gas flow expands through theturbine section to drive the compressor and the fan sections. Thecompressor section typically includes low and high pressure compressors,and the turbine section includes low and high pressure turbines.

An axial compressor includes a plurality of stages, each including tworows of airfoils—one rotating row, followed by one stationary row. Aclearance between the tip of each rotating airfoil and the housing orother static structure is controlled and minimized to improveefficiency. As the size of the compressor decreases the relative size ofthe clearance gap relative to both the blade span and chord, increases.A clearance to span and a clearance to chord ratios are larger forsmaller compressors. Larger clearance to chord ratios can contribute toinstability of the flow through the compressor airfoil row.

Turbine engine manufacturers continue to seek further improvements toengine performance including improvements to thermal, transfer andpropulsive efficiencies.

SUMMARY

In a featured embodiment, a compressor section for a gas turbine engineincludes a blade including a platform, a tip and an airfoil extendingbetween the platform and tip. The airfoil includes a root portionadjacent to the platform, a midspan portion and a tip portion. Each ofthe root portion, midspan portion and tip portion define a meridionalvelocity at stage exit with the tip portion including a first meridionalvelocity greater than a second meridional velocity of the midspanportion.

In another embodiment according to the previous embodiment, the tipportion includes between 65% and 100% of a radial span of the airfoiland the midspan portion includes between 20% and 65% of the radial spanof the airfoil.

In another embodiment according to any of the previous embodiments, thefirst meridional velocity produced at stage exit is between 6% and 15%greater than the second meridional velocity through the midspan portion.

In another embodiment according to any of the previous embodiments, afirst total pressure within the tip portion at stage exit is greaterthan a second total pressure across the midspan portion at stage exit.

In another embodiment according to any of the previous embodiments, thefirst total pressure at stage exit is between 1.2% and 3% greater thanthe second total pressure across the midspan portion at stage exit.

In another embodiment according to any of the previous embodiments, thecompressor section includes a plurality of stages and the blade includesa portion of at least one stage of the compressor section.

In another embodiment according to any of the previous embodiments, atleast one stage of the compressor section includes one of the last 5stages before an outlet of the compressor section.

In another embodiment according to any of the previous embodiments, thecompressor includes between 2 and 12 stages.

In another embodiment according to any of the previous embodiments, aclearance between the tip and a fixed structure of the compressorsection is greater than 1.5% of the radial span of the airfoil.

In another embodiment according to any of the previous embodiments, theclearance between the tip and a fixed structure of the compressorsection is between about 1.5% and 3% of the radial span of the airfoil.

In another featured embodiment, a blade for an axial compressor of a gasturbine engine includes a platform, a tip and an airfoil extendingbetween the platform and the tip. The airfoil includes a root portionadjacent the platform between 0% and 20% of a radial span of theairfoil, a midspan portion between 20% and 65% of the radial span of theairfoil and a tip portion between 65% and 100% of the radial span of theairfoil. Each of the root portion, midspan portion and tip portiondefine a meridional velocity at stage exit with the tip portionincluding a first meridional velocity greater than a second meridionalvelocity through the midspan portion.

In another embodiment according to the previous embodiment, the firstmeridional velocity produced at stage exit is between 6% and 15% greaterthan a second meridional velocity through the midspan portion.

In another embodiment according to any of the previous embodiments, afirst total pressure within the tip portion at stage exit is greaterthan a second total pressure across the midspan portion at stage exit.

In another embodiment according to any of the previous embodiments, thefirst total pressure at stage exit is between 1.2% and 3% greater than asecond total pressure across the midspan portion at stage exit.

Another embodiment, a method of operating a compressor section of a gasturbine engine includes configuring a blade to include an airfoil havinga platform, a tip and an airfoil between the platform and the tip,including a root portion adjacent the platform, a midspan portion and atip portion. A first meridional velocity is generated across the tipportion of the airfoil greater than a second meridional velocity acrossthe midspan portion.

Another embodiment according to the previous embodiment, includesgenerating the first meridional velocity with the tip portion that isbetween 6% and 15% greater than a second meridional velocity through themidspan portion.

In another embodiment according to any of the previous embodiments,includes generating a first total pressure with the tip portion that isgreater than a second total pressure across the midspan portion.

In another embodiment according to any of the previous embodiments, thegenerated first total pressure is between 1.2% and 3% greater than thegenerated second total pressure across the midspan portion.

Although the different examples have the specific components shown inthe illustrations, embodiments of this disclosure are not limited tothose particular combinations. It is possible to use some of thecomponents or features from one of the examples in combination withfeatures or components from another one of the examples.

These and other features disclosed herein can be best understood fromthe following specification and drawings, the following of which is abrief description.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a schematic view of an example gas turbine engine.

FIG. 2 is a schematic view of an example compressor section.

FIG. 3 is a schematic view of an example compressor blade.

FIG. 4 is a graph illustrating an example airfoil pressure profile.

FIG. 5 is a graph illustrating an example airfoil velocity profile.

FIG. 6 is a graph illustrating another example airfoil pressure profile.

FIG. 7 is a graph illustrating another airfoil velocity profile.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a core engine section 25 that includes a fansection 22, a compressor section 24, a combustor section 26 and aturbine section 28. Alternative engines might include an augmentorsection (not shown) among other systems or features. The fan section 22drives air along a bypass flow path B in a bypass duct defined within anacelle, while the compressor section 24 drives air along a core flowpath C for compression and communication into the combustor section 26then expansion through the turbine section 28. Although depicted as atwo-spool turbofan gas turbine engine in the disclosed non-limitingembodiment, it should be understood that the concepts described hereinare not limited to use with two-spool turbofans as the teachings may beapplied to other types of turbine engines including three-spoolarchitectures.

The exemplary engine 20 generally includes a low speed spool 30 and ahigh speed spool 32 mounted for rotation about an engine centrallongitudinal axis A relative to an engine static structure 36 viaseveral bearing systems 38. It should be understood that various bearingsystems 38 at various locations may alternatively or additionally beprovided, and the location of bearing systems 38 may be varied asappropriate to the application.

The low speed spool 30 generally includes an inner shaft 40 thatinterconnects a fan 42, a first (or low) pressure compressor 44 and afirst (or low) pressure turbine 46. The inner shaft 40 is connected tothe fan 42 through a speed change mechanism, which in exemplary gasturbine engine 20 is illustrated as a geared architecture 48 to drivethe fan 42 at a lower speed than the low speed spool 30. The high speedspool 32 includes an outer shaft 50 that interconnects a second (orhigh) pressure compressor 52 and a second (or high) pressure turbine 54.A combustor 56 is arranged in exemplary gas turbine 20 between the highpressure compressor 52 and the high pressure turbine 54. A mid-turbineframe 58 of the engine static structure 36 is arranged generally betweenthe high pressure turbine 54 and the low pressure turbine 46. Themid-turbine frame 58 further supports bearing systems 38 in the turbinesection 28. The inner shaft 40 and the outer shaft 50 are concentric androtate via bearing systems 38 about the engine central longitudinal axisA which is collinear with their longitudinal axes.

The core airflow is compressed by the low pressure compressor 44 thenthe high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded over the high pressure turbine 54 and lowpressure turbine 46. The mid-turbine frame 58 includes airfoils 60 whichare in the core airflow path C. The turbines 46, 54 rotationally drivethe respective low speed spool 30 and high speed spool 32 in response tothe expansion. It will be appreciated that each of the positions of thefan section 22, compressor section 24, combustor section 26, turbinesection 28, and fan drive gear system 48 may be varied. For example,gear system 48 may be located aft of combustor section 26 or even aft ofturbine section 28, and fan section 22 may be positioned forward or aftof the location of gear system 48.

The engine 20 in one example is a high-bypass geared aircraft engine. Ina further example, the engine 20 bypass ratio is greater than about six(6), with an example embodiment being greater than about ten (10), thegeared architecture 48 is an epicyclic gear train, such as a planetarygear system or other gear system, with a gear reduction ratio of greaterthan about 2.3 and the low pressure turbine 46 has a pressure ratio thatis greater than about five. In one disclosed embodiment, the engine 20bypass ratio is greater than about ten (10:1), the fan diameter issignificantly larger than that of the low pressure compressor 44, andthe low pressure turbine 46 has a pressure ratio that is greater thanabout five 5:1. Low pressure turbine 46 pressure ratio is pressuremeasured prior to inlet of low pressure turbine 46 as related to thepressure at the outlet of the low pressure turbine 46 prior to anexhaust nozzle. The geared architecture 48 may be an epicycle geartrain, such as a planetary gear system or other gear system, with a gearreduction ratio of greater than about 2.3:1. It should be understood,however, that the above parameters are only exemplary of one embodimentof a geared architecture engine and that the present invention isapplicable to other gas turbine engines including direct driveturbofans.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the engine 20 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet (10.67 km). The flight condition of 0.8 Mach and35,000 ft (10.67 km), with the engine at its best fuel consumption—alsoknown as “bucket cruise Thrust Specific Fuel Consumption (‘TSFC’)”—isthe industry standard parameter of lbm of fuel being burned divided bylbf of thrust the engine produces at that minimum point. “Low fanpressure ratio” is the pressure ratio across the fan blade alone,without a Fan Exit Guide Vane (“FEGV”) system. The low fan pressureratio as disclosed herein according to one non-limiting embodiment isless than about 1.45. “Low corrected fan tip speed” is the actual fantip speed in ft/sec divided by an industry standard temperaturecorrection of [(Tram° R)/(518.7° R)]0.5. The “Low corrected fan tipspeed” as disclosed herein according to one non-limiting embodiment isless than about 1150 ft/second (350 m/second).

The example gas turbine engine includes the fan section 22 thatcomprises in one non-limiting embodiment less than about twenty-six (26)fan blades 42. In another non-limiting embodiment, the fan section 22includes less than about twenty (20) fan blades. Moreover, in onedisclosed embodiment the low pressure turbine 46 includes no more thanabout six (6) turbine rotors schematically indicated at 34. In anothernon-limiting example embodiment the low pressure turbine 46 includesabout three (3) turbine rotors. A ratio between the number of fan blades42 and the number of low pressure turbine rotors is between about 3.3and about 8.6. The example low pressure turbine 46 provides the drivingpower to rotate the fan section 22 and therefore the relationshipbetween the number of turbine rotors 34 in the low pressure turbine 46and the number of blades 42 in the fan section 22 disclose an examplegas turbine engine 20 with increased power transfer efficiency.

Referring to FIG. 2 with continued reference to FIG. 1 , the highpressure compressor section 52 includes a plurality of stages 62 a-ibetween an inlet 74 and an outlet 72. Each stage 62 a-i includes arotating blade 66 a-i and a static vane 64 a-i. In this example thecompressor section 52 includes nine (9) stages. In another embodimentthe compressor section 52 may include between two (2) and twelve (12)stages. Moreover, it should be understood that although the disclosedcompressor section 52 embodiment includes nine (9) stages, a compressorwith more or less stages may benefit from this disclosure and is withinthe contemplation of this disclosure.

Clearance gaps 76, 78 between a tip of each compressor blade 66 a-i andthe housing 74 are present and minimized to increase engine operatingefficiency. The size of the gaps 76, 78 are commonly referred to inrelationship to a percent of a span of the airfoil. Accordingly, thegaps 76, 78 may be referred to as for example being between 1% and 2% ofthe overall airfoil span. There is a minimum size clearance between theblades 66 a-I and the housing 74 required to accommodate relativenon-axial movement caused by mechanical and thermal stress during engineoperation.

In the disclosed example, the clearance gap 76 is less than about 2% ofairfoil radial span. In other disclosed example, the clearance gap 76 isbetween about 0.5% and 0.8% of the airfoil radial span. The exampleclearance gap 78 ratio relative to the airfoil span is greater than thatof the clearance gap 76 for the larger blades near the inlet 70. In onedisclosed example the clearance gap 78 is greater than about 1.6%. Inanother example embodiment the clearance gap 78 is between about 2.5%and 3% of the airfoil radial span.

The radial length of each stage 62 a-i decreases in a direction from theinlet 70 toward the outlet 72. Gaps for the radially larger stages 62a-i as indicated by the gap 76 are smaller in relationship to the sizeof the radial airfoil span. As the radial span of the compressor stages62 a-i decreases, the gap becomes larger in relationship to the overallairfoil span as indicated schematically by gap 78. Disadvantageously,the ability to control clearances does not improve at the same rate thata diameter decreases. In other words, the clearance gap to airfoil spanratios increase as the radial span of each stage 62 a-I decreases.Accordingly, the gap 78 towards the outlet 72 is larger as a percent ofairfoil span compared to the gap 76 towards the inlet of the compressorsection 52.

Increases in gaps 76, 78 result in inefficiencies that not only affectthe one particular stage 62 a-i but can multiply across several stages.Airflows at the tip of the airfoil near the gaps slow relative toairflows radially inward across the airfoil. The slowing of airflow atthe airfoil tips generate unintended flow characteristics that propagatethrough subsequent stages 62 a-i. However, simply reducing the clearancegap is limited due to the need to accommodate both mechanical andthermal stresses encountered during engine operation. Accordingly,alternate methods and structures are needed to improve compressorefficiency.

Referring to FIG. 3 with continued reference to FIG. 2 , an examplecompressor blade 66 is illustrated and includes a platform 80, a tip 82and an airfoil 92 having a pressure side 84 and a suction side 86extending between the platform 80 and the tip 82. The airfoil 92includes a leading edge 88 and a trailing edge 90. A radial span of theairfoil 92 is schematically indicated at 106 and includes severalportions. In the disclosed embodiment, the airfoil 92 is divided into aroot portion 94, a midspan portion 96 and a tip portion 98. Asappreciated, the specific relationship and span ratios for each of theroot portion, mid-span portion and tip portion are disclosed by way ofexample and other divisions and sizes of each portion for other airfoilconfigurations are within the contemplation of this disclosure.

The example compressor blade 66 includes an airfoil 92 shape thatgenerates desired velocity and pressure profiles across the radial span106 between the platform 80 and the tip 82. The velocity and pressureprofiles are indicative of air flows across the airfoil 92 from theleading edge 88 toward the trailing edge 90. The gap clearances 76, 78influence the velocity and pressure profiles as well as the specificconfiguration of the airfoil 92.

An airfoil shape may be defined as series of x, y and z coordinatesalong the radial span. The specific airfoil shape is determined toprovide desired velocity and pressure profiles. Accordingly, thephysical geometry of the example blade 66 may be described and disclosedby specific velocity and pressure profiles.

Referring to FIG. 4 with continued reference to FIGS. 2 and 3 , a totalpressure profile 108 is shown including a total pressure 116 for a givenpercent of the airfoil span indicated 115. The disclosed span 115 beginsat 0% adjacent the platform 80 and extends radially outward to 100% spanat the tip 82. The total pressure 116 is a normalized pressure valueacross the airfoil at the radial span 115.

The example blade 66 is configured to generate the pressure profile 108with a bias toward an increase in pressure within the tip portion 98 ofthe radial span 106 compared to the midspan portion 96 as indicated at120. In the disclosed example, the tip portion 98 extends from 65% to100% of the radial span 106 of the airfoil 92. The midspan portionextends from between 20% and 65% of the radial span 106.

The disclosed blade 66 is configured to generate the disclosed pressureprofile 108 with the highest pressures biased within the tip portion 98of the radial span 106. In one disclosed example, the total pressure 116within the tip portion 98 is between about 1.2% and 3% greater than thetotal pressure 116 within the midspan portion 96 as indicated at 120. Inanother example embodiment, the total pressure in the tip portion 98 isbetween about 1.2% and 1.8% greater than the total pressure in themidspan portion 96.

Referring to FIG. 5 with continued reference to FIGS. 2-4 , the totalpressure profile 108 generates a desired airflow velocity profile 110along the radial span 106 of the blade 66. The airflow flow velocity 118is biased to have increased velocities toward the tip 82 to overcomeairflow degradation caused at the clearance gap 78. The disclosedvelocity 118 is the meridional velocity and is a measure of velocitywithin a plane defined along the engine axis A and the radial plane ofthe airfoil 92. The meridional velocity does not include thecircumferentially directed component of flow through the compressorsection 52.

The meridional velocity within the tip portion 98 is greater than themeridional velocity within the midspan portion 96. In this embodiment,the meridional velocity 118 within the tip portion 98 is greater thanthe meridional velocity within the midspan portion 96 by a quantityindicated at 122. In one disclosed embodiment, airflow velocities withinthe tip portion 98 are between 6% and 15% greater than velocities withinthe midspan portion 96. In another disclosed embodiment, velocitieswithin the tip portion 98 are between about 6% and 10% greater than thevelocities within the midspan portion 96.

In the example embodiment, at least one of the stages 62 a-I includes ablade 66 that provides the disclosed total pressure and velocityprofiles 108, 110. In this example the compressor section 52, theprofile 108 and velocity profile 110 are set up at the inlet to theeighth stage 62H by previous blade stages 62 a-G. The eighth stage 62Hthen propagates the total pressure profile 108 and velocity profile 110of the eight stage 62H to the subsequent stages. In this example, theninth stage 62I immediately downstream of the eighth stage 62H. Theninth stage 62I is the last stage of the compressor section 52 andbenefits from the airfoil configuration of the eighth stage 62H.

Referring to FIGS. 6 and 7 with continued reference to FIGS. 2-5 , FIG.6 illustrates a total pressure profile 112 for the ninth stage 62I. FIG.7 illustrates a velocity profile 114 of the ninth stage 62I. The ninthstage 62I of the compressor section does not include an airfoil with thesame configuration as the eight stage 62H but still benefits from thefavorable airflow velocities generated by blade 66. The total pressureprofile 112 of the ninth stage 62I includes an increased total pressureindicated at 124 within the tip portion 98 as compared to the midspanportion 96. The pressure profile 112 results in the correspondingvelocity profile 126 (FIG. 7 ). The velocity profile 114 includes anincrease, indicated at 126, in velocities within the tip portion ascompared to the midspan portion 96. The increases in total pressure 124,and airflow velocity 126 are a result of the airfoil configurationprovided by the eight stage blade 62H. Accordingly, the beneficialconfiguration provided by the tip biased velocities upstream of theblade 62H propagate downstream to further improve efficiencies in latercompressor stages.

Accordingly, the example airfoil configuration generates an increase invelocities near the tip portion of later stages of a compressor toinhibit and minimize air flow disturbances generated by increasedclearances.

Although an example embodiment has been disclosed, a worker of ordinaryskill in this art would recognize that certain modifications would comewithin the scope of this disclosure. For that reason, the followingclaims should be studied to determine the scope and content of thisdisclosure.

What is claimed is:
 1. A compressor section for a gas turbine enginecomprising: a plurality of blade stages including a first blade stageand a second blade stage immediately downstream of the first bladestage, wherein the first blade stage includes a blade including: a firstplatform; a first tip; a first leading edge and a first trailing edgeextending between the first platform and the first tip; and a firstairfoil extending between the first platform and the first tip, thefirst airfoil including a first root portion adjacent to the firstplatform, a first midspan portion and a first tip portion, a firstairfoil shape is configured to provide a first meridional velocity inthe second blade stage outside of a second midspan portion thatincreases relative to a second meridional velocity at 50% of a radialspan within a second midspan portion in a direction away from the secondmidspan portion in the second blade stage, wherein the first meridionalvelocity is between 6% and 10% greater than the second meridionalvelocity through the second midspan portion.
 2. The compressor sectionas recited in claim 1, wherein the second blade stage includes a secondairfoil shape different than the first airfoil shape.
 3. The compressorsection as recited in claim 1, wherein the first airfoil shape isdefined as a series of coordinates that provide a meridional velocityprofile across a second airfoil within the second blade stage withineach of a second root portion, the second midspan portion and a secondtip portion of airflow flowing across the second airfoil of the secondblade stage from a second leading edge to a second trailing edge duringoperation of the compressor section.
 4. The compressor section asrecited in claim 1, wherein a first total pressure within a second tipportion of the second blade stage is greater than a second totalpressure within the second midspan portion.
 5. The compressor section asrecited in claim 4, wherein the first total pressure is between 1.2% and1.8% greater than the second total pressure within the second midspanportion.
 6. The compressor section as recited in claim 4, wherein thefirst total pressure is between 1.2% and 3% greater than the secondtotal pressure within the second midspan portion.
 7. The compressorsection as recited in claim 1, wherein a clearance between each of thefirst tip and a second tip of a second airfoil within the second bladestage and a fixed structure of the compressor section is greater than1.5% of the radial span of respective ones of the first airfoil and thesecond airfoil.
 8. The compressor section as recited in claim 1, whereinthe clearance between the first tip and a second tip of a second airfoilwithin the second blade stage and a fixed structure of the compressorsection is between 1.5% and 3% of the radial span of the correspondingfirst airfoil and the second airfoil.
 9. The compressor section asrecited in claim 1, wherein the first meridional velocity is between 6%and 10% greater than the second meridional velocity through the secondmidspan portion.
 10. The compressor section as recited in claim 1,wherein the first blade stage and the second blade stage comprise atleast two of the last five blade stages of the plurality of stages ofthe compressor section before an outlet of the compressor section. 11.The compressor section as recited in claim 1, wherein the compressorsection comprises a high pressure compressor section disposed aft of alow pressure compressor section.
 12. The compressor section as recitedin claim 11, wherein the high pressure compressor is coupled to a highpressure turbine and the low pressure compressor section is coupled to alow pressure turbine disposed aft of the high pressure turbine.
 13. Acompressor section for a gas turbine engine comprising: a plurality ofblade stages including a first blade stage and a second blade stageimmediately downstream of the first blade stage, wherein the first bladestage includes a blade including: a first platform; a first tip; a firstleading edge and a first trailing edge extending between the firstplatform and the first tip, a first airfoil extending between the firstplatform and the first tip, the first airfoil including a first rootportion adjacent to the first platform, a first midspan portion and afirst tip portion, a first airfoil shape is configured to provide afirst meridional velocity in the second blade stage outside of a secondmidspan portion that increases relative to a second meridional velocityat 50% of a radial span within a second midspan portion in a directionaway from the second midspan portion in the second blade stage, whereinthe first airfoil shape is defined as a series of coordinates thatprovide a meridional velocity profile across a second airfoil within thesecond blade stage within each of a second root portion, the secondmidspan portion and a second tip portion of airflow flowing across thesecond airfoil of the second blade stage from a second leading edge to asecond trailing edge during operation of the compressor section andwherein each of the first tip portion and the second tip portioncomprises between 65% and 100% of a radial span of the corresponding oneof the first airfoil and the second airfoil and each of the firstmidspan portion and the second midspan portion comprises between 20% and65% of the radial span of the airfoil.
 14. A compressor section for agas turbine engine comprising: a plurality of blade stages including afirst blade stage and a second blade stage immediately downstream of thefirst blade stage, wherein the first blade stage includes a bladeincluding: a first platform; a first tip; a first leading edge and afirst trailing edge extending between the first platform and the firsttip; and a first airfoil extending between the first platform and thefirst tip, the first airfoil including a first root portion adjacent tothe first platform, a first midspan portion and a first tip portion, afirst airfoil shape is configured to provide a first meridional velocityin the second blade stage outside of a second midspan portion thatincreases relative to a second meridional velocity at 50% of a radialspan within a second midspan portion in a direction away from the secondmidspan portion in the second blade stage, wherein the plurality ofblade stages comprise nine blade stages and the first blade stage andthe second blade stage comprise the eighth blade stage and the ninthblade stage respectively.
 15. A compressor section for a gas turbineengine comprising: a plurality of blade stages including a first bladestage and a second blade stage immediately downstream of the first bladestage, wherein the first blade stage includes a first airfoil includinga first root portion adjacent to a first platform, a first midspanportion and a first tip portion; wherein the second blade stage includesa second airfoil including a second root portion adjacent to a secondplatform, a second midspan portion and second tip portion; wherein thefirst airfoil includes a shape that provides, during operation of thecompressor, a first meridional velocity in the second blade stageoutside of the second midspan portion is between 6% and 15% greater thana second meridional velocity within the second midspan portion.
 16. Thecompressor section as recited in claim 15, wherein the second meridionalvelocity is within between 40% and 55% of the radial span of the secondairfoil.
 17. The compressor section as recited in claim 15, wherein thefirst meridional velocity increases in a direction away from the secondmeridional velocity at between 45% and 55% of the radial span of thesecond airfoil.
 18. The compressor section as recited in claim 15,wherein a first total pressure within the second tip portion is greaterthan a second total pressure within the second midspan portion.
 19. Thecompressor section as recited in claim 18, wherein the first totalpressure is between 1.2% and 1.8% greater than the second total pressurewithin the second midspan portion.
 20. The compressor section as recitedin claim 18, wherein the first total pressure is between 1.2% and 3%greater than the second total pressure within the second midspanportion.